The present invention relates generally to gas turbine engines, and, more specifically, to turbines therein.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through multiple turbine stages. A turbine stage includes a stationary turbine nozzle having stator vanes which guide the combustion gases through a downstream row of turbine rotor blades extending radially outwardly from a supporting disk which is powered by extracting energy from the gases.
A first stage, or high pressure, turbine nozzle first receives the hottest combustion gases from the combustor which are directed to the first stage rotor blades which extract energy therefrom. A second stage turbine nozzle is disposed immediately downstream from the first stage blades, and is followed in turn by a row of second stage turbine rotor blades which extract additional energy from the combustion gases.
As energy is extracted from the combustion gases, the temperature thereof is correspondingly reduced. However, since the gas temperature is relatively high, the high pressure turbine stages are typically cooled by channeling through the hollow vane and blade airfoils cooling air bled from the compressor. Since the cooling air is diverted from the combustor, the overall efficiency of the engine is correspondingly reduced. It is therefore desired to minimize the use of such cooling air for maximizing overall efficiency of the engine.
The amount of cooling air required is dependent on the temperature of the combustion gases. That temperature varies from idle operation of the engine to maximum power operation thereof. Since combustion gas temperature directly affects the maximum stress experienced in the vanes and blades, the cooling air requirement for the turbine stages must be effective for withstanding the maximum combustion gas temperature operation of the engine although that running condition occurs for a relatively short time during engine operation.
For example, a commercial aircraft gas turbine engine which powers an aircraft in flight for carrying passengers or cargo experiences its hottest running condition during aircraft takeoff. For a military aircraft engine application, the hottest running condition depends on the military mission, but typically occurs during takeoff with operation of an afterburner. And, for a land-based gas turbine engine which powers an electrical generator, the hottest running condition typically occurs during the hot day peak power condition.
The maximum combustion gas temperature therefore varies temporally over the operating or running condition of the engine. And, the maximum combustion gas temperature also varies spatially both circumferentially and radially as the gases are discharged from the outlet annulus of the combustor. This spatial temperature variation is typically represented by combustor pattern and profile factors which are conventionally known.
Accordingly, each turbine stage, either blades or vanes, is typically specifically designed for withstanding the maximum combustion gas temperature experienced both temporarily and spatially in the combustion gases disposed directly upstream therefrom. Since the airfoils in each row of vanes and blades are identical to each other, the cooling configurations therefor are also identical and are effective for providing suitable cooling at the maximum combustion gas temperatures experienced by the individual stages for maintaining the maximum airfoil stress, including thermal stress, within acceptable limits for ensuring a suitable useful life of the turbine stages.
Furthermore, as engines wear during normal use in operation, combustion gas temperature may be intentionally increased within limits for ensuring minimum rated power for the engine notwithstanding deterioration thereof. Normal engine deterioration over extended use decreases its efficiency and resulting output power, with a loss in output power being regained by increasing the temperature of the combustion gases.
Accordingly, the turbine cooling configurations must be additionally effective for acceptable cooling in worn engines up to the typical exhaust gas temperature (EGT) limit.
It is therefore desired to provide a gas turbine engine turbine having improved cooling of the airfoils thereof.
A turbine includes three rows of airfoils which receive in sequence hot combustion gases during operation. The third row airfoils are clocked circumferentially relative to the first row airfoils for bathing the third row airfoils with relatively cool wakes discharged from the first row airfoils during the hottest running condition of the gas turbine engine being powered. The third row airfoils therefore avoid the hottest temperature of the combustion gases for reducing the cooling requirements thereof.